Thrust reverser mechanisms are commonly used in jet engines to provide additional braking power. Typically, in a thrust reverser mechanism at least a part of the normal flow of gases through the engine is diverted and forced to exit through side openings of the jet engine nacelle, rather than through the rear of the engine. Thus, when the thrust reverser mechanism is in use, gas flowing through the thrust reverser exits the engine generally perpendicular to the direction of normal flow that produces thrust to propel the aerospace vehicle forward. Vanes, alternatively called airfoils, positioned in the side openings, then direct the gas forwardly as it exits the engine nacelle, producing a reverse thrust, or a braking effect. Combinations of these vanes are referred to as cascades.
Known problems with thrust reversers are that the forwardly directed gas impinges upon the wing and fuselage of the aerospace vehicle, interferes with the proper operation of aerospace vehicle control surfaces, and is reingested by the engine. One attempted solution to this problem has been to produce individual cascade elements that not only redirect the gas flow forwardly, but also sideways in a predefined efflux pattern, that is, in a direction substantially tangential to the annular fan duct. Thus, when the thrust reverser is operating, the gas is directed forwardly in a designed efflux pattern that avoids impingement upon the wing, fuselage, and flight control surfaces, and that avoids reingestment by the engine.
The individual cascade elements are shaped as segments of a cylinder or a cone, and assembled concentrically around the centerline of the jet engine to form the completed cascade assembly. FIG. 1 shows a prior art cascade element 10 in the form of a generally rectangular cylindrical segment. Many such similar cascade elements are assembled to form a substantially cylindrically-shaped prior art cascade assembly 12 as shown in FIG. 2. Alternatively, the cascade elements are planar elements that are assembled around the centerline of the engine to substantially approximate a cylinder or cone.
One major problem with the solution attempted above is that many different cascade elements are produced and used in the completed cascade assembly. Typically, prior art cascade assemblies use sixteen or more cascade elements per engine to achieve the desired efflux pattern. In addition, each cascade element in prior art cascade assemblies requires a mirror image element for use on an engine mounted on the opposite side of the aerospace vehicle, doubling the required number of different cascade elements.
The manufacture of these many different cascade elements results in very high production costs. The present invention addresses the above problem by providing a cascade assembly composed of a reduced number of cascade element configurations that are substantially identical and used in many different locations in the cascade assembly. Additionally, in a first preferred embodiment, these cascade elements are supported in a manner which reduces the load carried by the cascade elements.
These design characteristics allow the construction of a lower cost cascade assembly requiring fewer different parts, and the option of construction techniques using lower strength and less expensive materials than have previously been possible. As a result, production costs are significantly reduced.